Modulated turbine vane cooling

ABSTRACT

A vane structure includes a baffle movably mounted within an aperture, the baffle movable to control a cooling flow between a first cooling cavity and a second cooling cavity.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly to a turbine vane cooling system.

In high performance gas generator applications, highly variable turbineconfigurations facilitate operations over a wide range of conditions.Traditional methods adjust cooling airflow gross flows to the entiremodule through a valve arrangement in the turbine cooling flow supplysystem upstream of the turbine module. While effective, this approachmay not accommodate stage cooling changes due to varying work splitsbetween turbine module stages or specific cooling requirements ofspecific components due to different life failure modes.

SUMMARY

A vane structure for a gas turbine engine according to one disclosednon-limiting embodiment of the present disclosure includes an airfoilsection which defines an aperture, a first cooling cavity incommunication with the aperture and a second cooling cavity incommunication with the aperture, and a baffle movably mounted within theaperture, the baffle movable to control a cooling flow between the firstcooling cavity and the second cooling cavity.

In a further embodiment of the foregoing embodiment, the aperture iscircular in cross-section. In the alternative or additionally thereto,in the foregoing embodiment the baffle is circular in cross section. Inthe alternative or additionally thereto, in the foregoing embodiment thebaffle is rotatable about an axis of rotation. In the alternative oradditionally thereto, in the foregoing embodiment the airfoil section isrotatable about an axis of rotation.

In a further embodiment of any of the foregoing embodiments, the airfoilsection is fixed.

In a further embodiment of any of the foregoing embodiments, the airfoilsection is rotatable.

In a further embodiment of any of the foregoing embodiments, the secondcooling cavity is a convection cooled cavity.

In a further embodiment of any of the foregoing embodiments, the vanstructure further includes a third cooling cavity which provides a rotorpurge feed.

A vane structure for a gas turbine engine according to another disclosednon-limiting embodiment of the present disclosure includes an airfoilsection which defines an aperture, the aperture circular incross-section, and a baffle movably mounted within the aperture, thebaffle circular in cross-section.

In a further embodiment of the foregoing embodiment, the vane structurefurther includes a first cooling cavity in communication with saidaperture and a second cooling cavity in communication with saidaperture.

In a further embodiment of any of the foregoing embodiments, the firstcooling cavity is a film cooled cavity. In the alternative oradditionally thereto, in the foregoing embodiment the second coolingcavity is a convection cooled cavity. In the alternative or additionallythereto, the foregoing embodiment includes a third cooling cavity whichprovides a rotor purge feed.

A method of communicating a cooling airflow according to anotherdisclosed non-limiting embodiment of the present disclosure includesrotating a baffle within an airfoil section of a turbine vane.

In a further embodiment of the foregoing embodiment, the method furtherincludes rotating the airfoil section of the turbine vane.

In a further embodiment of any of the foregoing embodiments, the methodfurther includes distributing a cooling airflow between a film cooledcavity and a convection cooled cavity.

In a further embodiment of any of the foregoing embodiments, the methodfurther includes distributing a cooling airflow 80%-20% between a filmcooled cavity and a convection cooled cavity with the baffle.

In a further embodiment of any of the foregoing embodiments, the methodfurther includes distributing a cooling airflow 20%-80% between a filmcooled cavity and a convection cooled cavity with the baffle.

In a further embodiment of any of the foregoing embodiments, the methodfurther includes blocking a cooling airflow 20%-80% between a filmcooled cavity and a convection cooled cavity with the baffle.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is a partial exploded view of a vane ring of one turbine stagewithin a turbine section of the gas turbine engine, the vane ring formedfrom a multiple of segments;

FIG. 3 is a side view of a turbine section;

FIG. 4 is a sectional view of a turbine vane taken along line A-A inFIG. 3;

FIG. 5 is a sectional view of the turbine vane of FIG. 4 illustratingvarious cooling cavities defined therein;

FIG. 6 is a perspective view of a turbine vane segment;

FIG. 7 is a sectional view of the turbine vane of FIG. 4 with a bafflein a first position;

FIG. 8 is a sectional view of the turbine vane of FIG. 4 with the bafflein a second position;

FIG. 9 is a sectional view of the turbine vane of FIG. 4 with the bafflein a third position;

FIG. 10 is a schematic view of a turbine vane with modulated coolingaccording to another disclosed non-limiting embodiment; and

FIG. 11 is a schematic view of a turbine vane with modulated coolingaccording to another disclosed non-limiting embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines such as a three-spool (plus fan) engine wherein anintermediate spool includes an intermediate pressure compressor (IPC)between the LPC and HPC and an intermediate pressure turbine (IPT)between the HPT and LPT.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 (“LPC”) and a lowpressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42through a geared architecture 48 to drive the fan 42 at a lower speedthan the low spool 30. An exemplary reduction transmission is anepicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A whichis collinear with their longitudinal axes.

Core airflow is compressed by the low pressure compressor 44 then thehigh pressure compressor 52, mixed with the fuel and burned in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 54, 46 rotationally drive therespective low spool 30 and high spool 32 in response to the expansion.

The main engine shafts 40, 50 are supported at a plurality of points bybearing structures 38 within the static structure 36. It should beunderstood that various bearing systems 38 at various locations mayalternatively or additionally be provided.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 48can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3:1, and in another example is greaterthan about 2.5:1. The geared turbofan enables operation of the low spool30 at higher speeds which can increase the operational efficiency of thelow pressure compressor 44 and low pressure turbine 46 and renderincreased pressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine 46 is pressuremeasured prior to the inlet of the low pressure turbine 46 as related tothe pressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 44, and the low pressure turbine 46has a pressure ratio that is greater than about five (5:1). It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentdisclosure is applicable to other gas turbine engines including directdrive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path B due to the high bypass ratio. The fan section 22 ofthe gas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the gas turbine engine 20 at its best fuelconsumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“T”/518.7)^(0.5) in which “T” represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, a turbine nozzle segment 70 includes anarcuate outer vane platform segment 72 and an arcuate inner vaneplatform segment 74 radially spaced apart from each other. The arcuateouter vane platform segment 72 may form a portion of an outer coreengine structure and the arcuate inner vane platform segment may form aportion of an inner core engine structure to at least partially definean annular turbine nozzle core airflow flow path (FIG. 3).

The circumferentially adjacent vane platform segments 72, 74 definesplit lines which thermally uncouple adjacent turbine nozzle segments 70which may be sealed therebetween, with, for example only, spline seals.That is, the temperature environment of the high pressure turbine 54 andthe substantial aerodynamic and thermal loads are accommodated by theplurality of circumferentially adjoining nozzle segments 70 whichcollectively form a full, annular ring about the centerline axis A ofthe engine.

Each turbine nozzle segment 70 may include a multiple (two shown) ofcircumferentially spaced apart airfoil sections 78 which extend radiallybetween the vane platform segments 72, 74. It should be appreciated theany number of vanes may define each segment. Alternatively, the vanesmay be formed as a unitary full, annular ring.

With reference to FIG. 3, the high pressure turbine 54 generallyincludes a turbine case 60 with a multiple of turbine stages. The highpressure turbine 54 includes a multiple of rotor structures (two shown;62A, 62B) interspersed with a vane structure (one shown; 64A). Each ofthe rotor structures 62A, 62B and the vane structure 64A, includerespective airfoil sections 76, 78. It should be appreciated that anynumber of stages will benefit herefrom and although schematicallydepicted as the high pressure turbine 54 in the disclosed embodiment, itshould also be understood that the concepts described herein are notlimited to use with high pressure turbines as the teachings may beapplied to other sections such as low pressure turbines, power turbines,intermediate pressure turbines as well as other cooled airfoilstructures with any number of stages.

In one disclosed non-limiting embodiment, a baffle 84 is movably mountedthrough each vane airfoil section 78 and the respective outer vaneplatform segment 72. The baffle 84 is in fluid communication with asecondary airflow source S which originates from the compressor section24—illustrated schematically from compartment 86. As defined herein thesecondary airflow may be any relatively cooler airflow different than acore airflow C. The baffle 84 may terminate in the respective inner vaneplatform segment 74 to provide support and block a distal end of thebaffle 84 such that the secondary airflow is communicated through baffleholes 88 along the length thereof.

With reference to FIG. 4, each vane airfoil section 78 is defined by anouter airfoil wall surface 90 between the leading edge 92 and a trailingedge 94. The outer airfoil wall surface 90 is typically shaped for usein a respective stage of the high pressure turbine section. The outerairfoil wall surface 90 defines a generally concave shaped portionforming a pressure side 90P and a generally convex shaped portionforming a suction side 90S.

The baffle 84 is located within an aperture 96 which is in fluidcommunication with a first cooling cavity 98, a second cooling cavity100 and a third cooling cavity 102 (FIG. 5). The aperture 96 in thedisclosed non-limiting embodiment is circular in cross-section toreceive the baffle 84 for rotation about axis W. It should beappreciated that other movements of the baffle 84 such as axial slidingalong axis W may alternatively or additionally be provided. It should befurther appreciated that the cooling cavities 98, 100, 102 may be ofvarious shapes and sizes to communicate the cooling airflow through theouter airfoil wall surface 90.

In the disclosed non-limiting embodiment, the first cooling cavity 98defines film cooling cavities; the second cooling cavity 100 definesconvective cooling and rotor purge feed cavities; and the third coolingcavity 102 defines convective cooling cavities. Additional oralternative cavities may also benefit herefrom. The rotor purge feedcavities of the second cooling cavity 100 communicate cooling airflowthrough the vane airfoil section 78, the respective inner platformsegment 74 and into the downstream rotor 62B (illustrated schematicallyby arrows Rs); FIG. 3.

With reference to FIG. 6, an actuator system 104 such as a unison ring(illustrated schematically) rotates an actuator arm 106 and therebyrotates the baffle 84. The actuator system 104 rotates each baffle 84within the aperture 96 to change the position of the baffle holes 88 toselectively vary cooling modulation from a relatively high flow filmcooled scheme (FIG. 7) to a relatively high convectively cooled scheme(FIG. 8). The baffle 84 may also be rotated to essentially minimize oreliminate cooling flow altogether (FIG. 9).

In the disclosed, non-limiting embodiment, the relatively high flow filmcooled scheme (FIG. 7) may provide approximately 80% film cooling and20% convective cooling while the relatively high convectively cooledscheme (FIG. 8) may provide 20% film cooling and 80% convective cooling.It should be appreciated that various arrangements of baffle holes 88may provide various cooling schemes. The baffle 84 can be articulatedsuch that the baffle holes 88 will deliver primarily blade cooling air,primarily vane cooling air, both vane and blade cooling air, or nocooling air. The ability to modulate cooling airflow facilitates customtailored parasitic flow to each operating point with specificationtoward a single component row or collection of rows. That is, the baffle84 provides a local method of modulating cooling airflow.

The angular position of the baffle 84 relative to axis W dictates whichcavities 98, 100, 102 the baffle holes 88 supply. The amount of airdrawn out of the baffle 84 may be dependent on the exit conditions ofthe respective cavity 98, 100, 102, whether the cavity supplies amultiplicity of film holes (FIG. 7) or if the baffle 84 is positionedsuch that cooling airflow is blocked or directed to low flow areas (FIG.9).

With reference to FIG. 10, another disclosed non-limiting embodimentutilizes a rotationally fixed baffle 84′ upon which the vane airfoilsection 78′ may pivot. That is, the vane airfoil section 78′ is avariable vane within which the cooling airflow is selectivelycommunicated to various cavities dependent upon the position of the vaneairfoil portion 78′ with respect to the baffle 84′.

With reference to FIG. 11, another disclosed non-limiting embodimentutilizes a rotational baffle 84″ and a rotational vane airfoil section78″. That is, the vane airfoil section 78″ is a variable vane and thebaffle 84″ rotates as described above. The rotational baffle 84″ and therotational vane airfoil section 78″ may be actuated through respectiveunison ring to provide significant controllability such that coolingairflow is selectively communicated to various cavities at variousrotational positions of the vane airfoil section 78″.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” “bottom”, “top”,and the like are with reference to the normal operational attitude ofthe vehicle and should not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A vane structure for a gas turbine enginecomprising: an airfoil section which defines an aperture, said aperturecircular in cross-section; a baffle movably mounted within saidaperture, said baffle circular in cross-section and configured to rotateabout an axis; a first cooling cavity in communication with saidaperture; a second cooling cavity in communication with said aperture;and a third cooling cavity in communication with said aperture, whereinsaid first cooling cavity defines film cooling cavities, wherein saidsecond cooling cavity defines convective cooling and rotor purge feedcavities, and wherein said third cooling cavity defines convectivecooling cavities.
 2. The vane structure as recited in claim 1, whereinsaid baffle includes a plurality of baffle holes along a length of thebaffle and terminates in a vane platform segment to block a distal endof the baffle such that a cooling airflow is configured to becommunicated through the plurality of baffle holes.
 3. The vanestructure as recited in claim 1, further comprising: a vane platformsegment, wherein said baffle terminates in the vane platform segment andincludes a plurality of baffle holes along a length of the baffle.
 4. Amethod of communicating a cooling airflow comprising: rotating a bafflewithin an aperture of an airfoil section of a turbine vane about anaxis, wherein the turbine vane includes: a first cooling cavity incommunication with said aperture; a second cooling cavity incommunication with said aperture; and a third cooling cavity incommunication with said aperture, wherein said first cooling cavitydefines film cooling cavities, wherein said second cooling cavitydefines convective cooling and rotor purge feed cavities, and whereinsaid third cooling cavity defines convective cooling cavities.
 5. Themethod as recited in claim 4, further comprising: rotating the airfoilsection of the turbine vane.
 6. The method as recited in claim 4,further comprising: distributing a cooling airflow between the filmcooling cavities and the convective cooling cavities.
 7. The method asrecited in claim 4, further comprising: distributing a cooling airflow80%-20% between the film cooling cavities and the convective coolingcavities with the baffle.
 8. The method as recited in claim 4, furthercomprising: distributing a cooling airflow 20%-80% between the filmcooling cavities and the convection cooling cavities with the baffle. 9.The method as recited in claim 4, further comprising: blocking a coolingairflow 20%-80% between the film cooling cavities and the convectioncooling cavities with the baffle.
 10. The method as recited in claim 4,wherein said baffle terminates in a vane platform segment and includes aplurality of baffle holes along a length of the baffle.